1. Field of the Invention
The present invention relates to an axial compressor as a main component of an industrial use gas turbine, a jet engine, etc.
2. Description of the Prior Art
In an industrial use gas turbine or a jet engine, a gas turbine is rotated for obtaining an output by a fuel combustion at a combustor using a compressed air.
At this time, at an axial compressor operating for compressing air, in order to alleviate a load acting per blade row, blade rows are increased in number and thus a predetermined output and efficiency is obtained. For this reason, while a certain efficiency is secured, the apparatus becomes larger unavoidably.
That is, as shown by one example of a stationary blade profile in the prior art of FIG. 5, since a stationary blade 1 which is fixed to an end wall face 3 is provided straight-uprightly so that a projection of any blade cross section is overlapped at one position, when the air flowing from an inlet of the compressor as shown by an arrow A makes contact with the end wall face 3 formed by a shroud etc. and the other end wall face (not shown) formed by a casing etc. (not shown), there is structurally occurring and growing a boundary layer flow 4 at a respective vicinity of said end wall faces.
Such occurring and growing of the boundary layer flow as mentioned above bring about a loss at the vicinity of the blade end wall faces and a lowering of efficiency and yet that becomes larger as the load per blade row increases, thus there is such a shortcoming that, in order to obtain a compressed air of a predetermined pressure efficiently by an industrial use gas turbine etc., the load per blade row must be reduced and the blade rows must be increased in number by so much, thus the entire apparatus becomes larger unavoidably.